Fuel mixing in a gas turbine engine

ABSTRACT

A gas turbine engine has a device for admitting a rotating fluid flow from an annular space of the casing to the inlet portion of a combustor to form a rotating fluid flow in the inlet portion of the combustor. The rotating fluid flow is formed in the annular space of the casing by supplying a fluid from a compressor to the blades of the turbine rotor disk.

[0001] This application cross-references three copending U.S. patentapplications, each of which was filed on Sep. 25, 1998, as U.S. patentapplication Ser. No. 09/______, 09/______ and 09/______, respectively,each of which copending U.S. applications is incorporated herein byreference.

[0002] The invention relates to the field of gas turbine engines, andmore specifically, to an improved gas turbine engine using a rotatingfluid flow train to feed the combustor and enhance air/fuel mixing andemissions.

BACKGROUND OF THE INVENTION

[0003] A type of prior art gas turbine engine has a compressor, a fuelsource, a combustion air source, a casing, and a combustor to prepare aheated fluid from fuel and combustion air. The combustor is connected tothe fuel source, the combustion air source and the compressor.Practically the entire fluid flow from the compressor is directed to thecombustor. The engine has a turbine rotor disk with blades that receivethe heated fluid from the combustor. As the turbine rotor disk rotatesduring engine operation, the heated fluid flow coming from the combustorhas to be directed at an angle to the blades to ensure smooth entryconditions. This is done using stator vanes that are positioned at acertain angle and direct the heated fluid from the combustor to theturbine rotor disk in a manner compatible with rotor disk rotation. Thisgas turbine engine is disclosed in U.S. Pat. No. 3,826,084 to Branstromet al.

[0004] The stator vane angle normally is chosen to accommodate the mostoptimum and prevailing turbine rotor disk operating conditions (speed).This solution is quite acceptable for gas turbine engines that have moreor less stable operating conditions, such as when used for powergeneration. In applications where the load upon the gas turbine engineis steady, the turbine rotor disk rotates at a stable speed, and theentry angle for the blades remains unchanged thus minimizing losses. If,on the other hand, this gas turbine engine is used to power a vehicle,the situation is radically different. In that application, the turbinerotor disk speed will vary within a broad range depending on vehicleload. Consequently, the entry angle also varies within a broad rangeunder load fluctuations, which leads to greater losses. This problemcould not be solved by using the conventional approach with the statorvanes. It is possible to use controllable stator vanes to change theentry angle at the blades, but it is a very complicated and expensivesolution given the high temperatures downstream of the combustor andspace limitations. As a result, the gas turbine engine would have highlosses in vehicle applications. Moreover, the stator and vanes occupiesan additional space and makes the engine design more complicated andexpensive. The use of controllable vanes makes the engine less reliable.

[0005] The problems indicated above are solved in the gas turbine engineof this invention.

SUMMARY OF THE INVENTION

[0006] It is an object of the invention to provide a gas turbine engineof the above type that has a higher efficiency.

[0007] Another object of the invention is to provide a more compact gasturbine engine that has a simpler design.

[0008] Another object of the invention is to improve the emissioncharacteristics of the gas turbine engine.

[0009] A gas turbine engine has a device to admit a rotating fluid flowfrom an annular space in the casing to the inlet portion of a combustorto form a rotating fluid flow in the inlet portion of the combustor. Therotating fluid flow is formed in the annular space of the casing bysupplying a fluid from a compressor to the blades of the turbine rotordisk.

[0010] Other objects and advantages of the invention will becomeapparent from the following detailed description of preferredembodiments and accompanying drawings.

DETAILED DESCRIPTION OF THE DRAWINGS

[0011]FIG. 1 shows a diagrammatic view of a gas turbine engine accordingto the invention.

[0012]FIG. 2 is a sectional view of an embodiment of the annular space(Leonid to supply a sketch.)

DETAILED DESCRIPTION OF THE DRAWINGS

[0013] With reference to FIG. 1, a gas turbine engine has a casing 10, acompressor 12 for supplying a compressed fluid, a turbine rotor disk 14mounted downstream of compressor 12 installed on the turbine rotor, acombustor 16 to prepare a heated fluid to be supplied to turbine rotordisk 14. Combustor 16 has a port 18 to admit fuel supplied from a fuelsource (not shown). Combustor 16 defines a combustion zone 20 in whichthe heated fluid is formed. Combustion air is supplied from an airsource (not shown) as shown by arrows A to an inlet portion of thecombustor in which port 18 is provided.

[0014] The inlet portion of the combustor shown at 19 is defined by aninner annular wall 22 of combustor 16 and by an annular guide wall 24that extends within the combustor in a spaced relation to annular innerwall 22. Annular guide wall 24 is installed by brackets 26 in such amanner that a space 28 is left for fluid passage.

[0015] A part of the fluid from compressor 12 is supplied to turbinerotor disk 14, bypassing combustor 16, as shown by arrows B, throughpassage 30 in casing 10 and reaching a zone 32 upstream of turbine rotordisk 14. Vanes 34 can be provided in passage 30 to make this fluid flowcompatible with the turbine rotor disk 14 rotation. These vanes willfunction in an optimum manner only under certain turbine engineoperating conditions. Since the quantity of fluid that is fed to theturbine rotor disk 14 and the velocity of this fluid are not very high,losses that would occur under non-optimum conditions would be relativelylow. This fluid is admitted to turbine rotor disk 14 and envelops theblades 15. The fluid from the compressor 12 passes through a passage 36of the blade 15 and leaves the passage 36 to reach an annular space 38that is defined in casing 10 and surrounds blades 15. As the blades 15rotate, the fluid from the compressor 12 leaves blade passage 36 havingobtained a rotation that forms a rotating fluid flow in annular space38. This rotating fluid flow is admitted through space 28 to inletportion 19 of combustor 16 to form a rotating fluid flow there. As fuelis fed through port 18, it is entrained in a rotary motion by therotating fluid flow in the inlet portion, and intense stirring andmixing of fuel and fluid will take place to prepare a good quality fuelmixture. The rotating fluid flow entrains air that is fed as shown byarrow A, moves into combustion zone 20, and imparts a spin to the heatedfluid when it is formed in combustion zone 20. The direction of thisrotating flow is the same as the turbine rotor disk direction ofrotation and the velocity of this rotating flow steadily follows turbinerotor disk 14 rotation velocity (with a very short lag). The heatedfluid formed in combustor 16 will move to the turbine blades 15 in amanner that is almost entirely compatible with rotation of the turbinerotor disk. Consequently, losses in this zone, which account for most ofthe losses in the turbine flow duct, are minimized.

[0016] Another advantage of the invention is that the fluid from thecompressor that goes through passage 36 and reaches blade 15 cools theblade and the adjacent wall of casing 10.

[0017] The intensive mixing and stirring of fuel, air, and the fluidthat comes from the compressor in inlet portion 19 provides almost idealconditions to prepare a fuel mixture. This high quality fuel mixtureprovides better conditions for combustion and improves the emissioncharacteristics of the engine.

[0018] Another advantage of the invention is the method of preparationof the fuel mixture. The quantity of fuel supplied for small-power gasturbine engines is rather low. It is very difficult to prepare ahomogeneous fuel mixture with a ratio of fuel to air and fluid of 1:15to 1:30. The fuel mixing method that is used here solves this problem.When fuel is entrained in a rotary motion by the rotating fluid flowadmitted to the inlet portion of the combustor, fuel atomizing, mixingand stirring in the rotating flow are very thorough and intensive. Thisthoroughness assures a high degree of homogeneity of the fuel mixture.

[0019]FIG. 2 shows an embodiment of the space 28 with annular guidewalls 24 attached by brackets 26. This space 28 can take the form of anarc slit cut in a flanged portion of the annular guide wall or in theform of spaces between the adjacent brackets (not shown).

I claim:
 1. A method of operation of a gas turbine engine having acompressor for producing a fluid flow, a casing, a combustor in saidcasing, said combustor having an inlet portion, a turbine rotor diskwith blades, and an annular space in said casing, said annular spacesurrounding said blades, said method comprising: supplying fuel andcombustion air to said combustor to prepare a heated fluid; supplyingsaid heated fluid directly from said combustor to said blades; supplyingsaid fluid flow from said compressor to said blades to form a rotatingfluid flow in said annular space; feeding at least a part of saidrotating fluid flow into said inlet portion of said combustor.
 2. Themethod of claim 1, wherein said fuel is supplied into said rotatingfluid flow within said inlet portion of said combustor.
 3. A method ofoperation of a gas turbine engine having a compressor for producing afluid flow, a casing, a combustor in said casing, said combustor havingan inlet portion, a turbine rotor disk with blades, and an annular spacein said casing, said annular space surrounding said blades, said methodcomprising: supplying said fluid flow from said compressor to saidblades to form a rotating fluid flow in said annular space; feeding atleast a part of said rotating fluid flow into said inlet portion of saidcombustor; feeding said fuel into said rotating fluid flow within saidinlet portion of said combustor preparing a heated fluid in saidcombustor by burning said fuel and air in said combustor; supplying saidheated fluid directly from said combustor to said blades.
 4. A gasturbine engine, said gas turbine engine comprising: a compressor forproducing a fluid flow; a fuel source; a combustion air source; acasing; a combustor in said casing, said combustor having an annularinner wall and an inlet portion, said combustor communicating with saidfuel source and with said combustion air source to prepare a heatedfluid; a turbine rotor disk with blades said blades positionedimmediately downstream of said combustor for receiving said heated fluidfrom said combustor; an annular space in said casing, said annular spacesurrounding said blades; a zone upstream of said turbine rotor disk,said zone communicating with said compressor for supplying said fluidflow from said compressor to said blades to form a rotating fluid flowin said annular space; a means for admitting said rotating fluid flowfrom said annular space to said inlet portion of said combustor, wherebya rotating fluid flow is formed in said inlet portion of said combustor.5. The gas turbine engine of claim 4, wherein said fuel sourcecommunicates with said inlet portion of said combustor.
 6. The gasturbine engine of claim 4, wherein said means for admitting saidrotating fluid flow from said annular space to said inlet portion ofsaid combustor comprises: an annular guide wall that is installed insaid combustor in a spaced relation to said annular inner wall of saidcombustor, said annular guide wall defining with said annular inner wallof said combustor said inlet portion of said combustor; said inletportion of said combustor communicating with said annular space of saidcasing.
 7. The gas turbine engine of claim 6, wherein said fuel sourcecommunicates with said inlet portion of said combustor.
 8. A gas turbineengine, said gas turbine engine comprising: a compressor for producing afluid flow; a fuel source; a combustion air source; a casing; acombustor in said casing, said combustor having an annular inner walland an inlet portion, said combustor communicating with said combustionair source to prepare a heated fluid by burning said fuel with saidcombustion air; a turbine rotor disk with blades said blades positionedimmediately downstream of said combustor for receiving said heated fluidfrom said combustor; an annular space in said casing, said annular spacesurrounding said blades; a zone upstream of said turbine rotor disk,said zone communicating with said compressor for supplying said fluidflow from said compressor to said blades to form a rotating fluid flowin said annular space; an annular guide wall that is installed in saidcombustor in a spaced relation to said annular inner wall of saidcombustor, said annular guide wall defining with said annular inner wallof said combustor said inlet portion of said combustor; said inletportion of said combustor communicating with said annular space of saidcasing; said fuel source communicates with said inlet portion of saidcombustor.